Most relatively small missiles in use today are propelled by solid fuel rockets as opposed to, for example, turbojet engines The selection of a solid fuel rocket as a propulsion device has been largely dictated by two factors. First, in many instances, a turbine engine cannot be fabricated sufficiently economically as to compete with a solid fuel rocket engine. Secondly, in small size missiles, i.e., those having relatively small diameter on the order of about six inches, it has heretofore been quite difficult to manufacture an efficient turbojet engine. The difficulty lies in the fact that the turbine jet engine must fit within the six inch envelope required of the propulsion unit for such a missile. Unfortunately, the use of solid fuel rocket engines has had consequences that are not desirable in many applications.
Specifically, the use of solid fuel rocket engines results in the loss of some degree of control of the missile flight path or trajectory. In contrast, control is far greater with gas turbine engines whose output can readily be varied. Further, even if the gas turbine engine operates relatively inefficiently, the use of such an engine would greatly extend the range of the missile.
The difficulty in economically producing small diameter gas turbine engines resides primarily in the labor intensive nature of the manufacture of the combustor. Furthermore, as combustor sizes shrink to fit within some desired envelope, the difficulty in achieving efficient combustion of fuel rises significantly In particular, as the size or volume of a combustor is reduced, there may be insufficient volume to allow the fuel to first be vaporized completely, burned efficiently, and then mixed uniformly.
In order to overcome the foregoing, a unique low cost annular combustor was developed as disclosed in commonly owned U.S. Pat. No. 4,794,754, issued Jan. 3, 1989. This annular combustor has proven to be well suited for its intended purpose, but it was desired to attempt to achieve greater thrust with a higher turbine inlet temperature while meeting the necessary size constraints and achieving the goal of ultralow cost for throw away missile applications. For this purpose, it was recognized that a new approach would be required to reach the necessary parameters of operation.
More specifically, the small missile application may typically be such as to require a spacing of one inch between combustor walls. It is also typically a necessary parameter that very difficult to burn missile fuel such as JP10 be handled efficiently even though such fuel is known to have a high carbon content together with very high surface tension and viscosity which, respectively, results in carbon buildup on the walls of the combustor together with a smoke filled exhaust and makes fuel atomization difficult which creates combustion inefficiency and flame instability problems. In very small combustors, such problems are oftentimes greatly magnified particularly where only low cost solutions are available.
In addition, such problems are exacerbated where high turbine inlet temperatures are to be found. This follows because only low cost means of cooling the metal at such high temperatures are acceptable especially for small missile applications. In addition, the turbine inlet temperature distribution must be exceptionally uniform to avoid burnout of the turbine nozzle blades.
The present invention is directed to overcoming one or more of the foregoing problems and achieving the resulting objects.